The present invention relates to combustion chambers for aircraft gas turbine engines, more particularly such combustion chambers having a deflecting feeder bowl assembly.
Gas turbine engines are known in which the annular combustion chamber is bound by inner and outer walls connected by a forward end wall, the forward end wall having openings through which fuel and air are injected into the combustion chamber via an air and fuel injection device. Typically, a feeder bowl structure is interposed between the air and fuel injection device and the forward end wall of the combustion chamber.
In such known combustion chambers, particularly those used in turbojet engines in military aircraft, the feeder bowl structures are subjected to high stresses. In high performance turbojet engines, the air is being injected into the combustion chambers at higher compression ratios, thereby increasing the pressures and temperatures at the intakes of the combustion chambers. The increases in pressure and temperature also relate to an increase in the gas temperatures within the combustion chamber, thereby subjecting the feeder bowls to high thermal and mechanical stresses. Typically, in modern turbojet engines, the feeder bowl structures are subjected to temperatures between 900.degree. and 1,000.degree. C. The feeder bowls are cooled by only approximately 2% of the air flow into the combustion chamber.
Such extreme operating conditions may cause burns and/or deformations in the outer edges of the feeder bowl, thereby deleteriously affecting the cooling film formed on the walls of the combustion chamber by the outer portions of the feeder bowls. The improper formation of the cooling film may result in deformation or complete burning through the wall of the combustion chamber.